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AIRCRAFT PARAMETRIC 3D MODELLING AND PANEL CODE ANALYSIS FOR CONCEPTUAL DESIGN Mehdi Tarkian Francisco Javier Zaldivar Tessier Division of Machine Design Degree Project Department of Management and Engineering LIU-IEI-TEK-A--07/00086--SE

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Page 1: AIRCRAFT PARAMETRIC 3D MODELLING AND PANEL CODE

AIRCRAFT PARAMETRIC 3D MODELLING AND PANEL CODE ANALYSIS FOR CONCEPTUAL DESIGN

Mehdi Tarkian Francisco Javier Zaldivar Tessier

Division of Machine Design

Degree Project Department of Management and Engineering

LIU-IEI-TEK-A--07/00086--SE

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Framläggningsdatum

2007/02/28 Publiceringsdatum (elektronisk version)

Institution och avdelning Institutionen för konstruktions- och produktionsteknik

Nyckelord CATIA, PANAIR, Panel Code Method, Parametric Aircraft Model, Aircraft design process

Sammanfattning Throughout the development of this report there will be a brief explanation of what the actual Aircraft Design Process is and in which stages the methodology that the authors are proposing will be implemented as well as the tools that will interact to produce this methodology. The proposed tool will be the first part of a methodology that, according to the authors, by integrating separate tools that are currently used in different stages of the aeronautical design, will promote a decrease in the time frame for the initial stages of the design process. The first part of the methodology above, that is proposed in this project, starts by creating a computer generated aircraft model and analyzing its basic aerodynamic characteristics “Lift Coefficient” and “Induced Drag Coefficient”, this step will be an alternative to statistical and empirical methods used in the industry, which require vast amount of data This task will be done in several steps, which will transfer the parametric aircraft model to an input file for the aerodynamic analysis program. To transfer the data a “translation” program has been developed that arranges the geometry and prepares the input file for analysis. During the course of this report the reader will find references to existing aircrafts, such as the MD-11 or Airbus 320. However, these references are not intended to be an exact computer model of the mentioned airplanes. The authors are using this as reference so the reader can relate what he/she is seeing in this paper to existing aircrafts. By doing such comparison, the author intends to demonstrate that the Parametric Model that has been created possesses the capability to simulate to some extend the shape of existing aircrafts. Finally from the results of this project it is concluded that the methodology in question is promising. Linking the two programs is possible and the aerodynamic characteristics of the models tested fall in the appropriate range. Non the less the research most continue following the line that has been discussed in this report

Titel: AIRCRAFT PARAMETRIC 3D MODELLING AND PANEL CODE ANALYSIS FOR CONCEPTUAL DESIGN Författare: Mehdi Tarkian Francisco Javier Zaldivar Tessier

URL för elektronisk version

Språk

Svenska X Engelska

Annat (ange nedan)

Rapporttyp

Licentiatavhandling X Examensarbete

C-uppsats D-uppsats Övrig rapport

ISBN: ISRN: LIU-IEI-TEK-A--07/0086--SE

Serietitel Serienummer/ISSN

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ABSTRACT Throughout the development of this report there will be a brief explanation of what the actual Aircraft Design Process is and in which stages the methodology that the authors are proposing will be implemented as well as the tools that will interact to produce this methodology. The proposed tool will be the first part of a methodology that, according to the authors, by integrating separate tools that are currently used in different stages of the aeronautical design, will promote a decrease in the time frame for the initial stages of the design process. The first part of the methodology above, that is proposed in this project, starts by creating a computer generated aircraft model and analyzing its basic aerodynamic characteristics “Lift Coefficient” and “Induced Drag Coefficient”, this step will be an alternative to statistical and empirical methods used in the industry, which require vast amount of data. This task will be done in several steps, which will transfer the parametric aircraft model to an input file for the aerodynamic analysis program. To transfer the data a “translation” program has been developed that arranges the geometry and prepares the input file for analysis. During the course of this report the reader will find references to existing aircrafts, such as the MD-11 or Airbus 310. However, these references are not intended to be an exact computer model of the mentioned airplanes. The authors are using this as reference so the reader can relate what he/she is seeing in this paper to existing aircrafts. By doing such comparison, the author intends to demonstrate that the Parametric Model that has been created possesses the capability to simulate to some extend the shape of existing aircrafts. Finally from the results of this project it is concluded that the methodology in question is promising. Linking the two programs is possible and the aerodynamic characteristics of the models tested fall in the appropriate range. None the less the research must continue following the line that has been discussed in this report.

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NOTATION .................................................................................................................................................................. 1

1 INTRODUCTION ..................................................................................................................................................... 4 1.1 BACKGROUND....................................................................................................................................................... 5

1.1.1 Conventional Design Process ...................................................................................................................... 6 1.1.2 Objective of this thesis ................................................................................................................................. 9 1.1.3 Planning..................................................................................................................................................... 11

2 PANAIR 502I ........................................................................................................................................................... 12 2.1 FUNDAMENTALS ................................................................................................................................................. 13

2.1.1 What is PANAIR......................................................................................................................................... 13 2.2 IMPLEMENTATION............................................................................................................................................... 14

2.2.1 Working with PANAIR ............................................................................................................................... 14 2.2.2 Input File.................................................................................................................................................... 16 2.2.3 Output File ................................................................................................................................................. 24

2.3 PANAIR SURFACE VALIDATION ........................................................................................................................ 28 2.3.1 Validation procedure ................................................................................................................................. 28 2.3.2 Grid Point Extraction................................................................................................................................. 30 2.3.3 Result comparison...................................................................................................................................... 33

3 CATIA V5 ................................................................................................................................................................ 36 3.1 PARAMETRIC AIRCRAFT MODEL......................................................................................................................... 37

3.1.1 About CATIA in this thesis ......................................................................................................................... 37 3.1.2 Building structure....................................................................................................................................... 38 3.1.3 Fuselage Part ............................................................................................................................................. 39 3.1.4 Wing Parts.................................................................................................................................................. 41 3.1.5 Horizontal Tail ........................................................................................................................................... 43 3.1.6 Vertical Tail ............................................................................................................................................... 44 3.1.7 Engine Configuration................................................................................................................................. 45 3.1.8 Engines choice and positioning.................................................................................................................. 46

3.2 PARAMETERIZATION........................................................................................................................................... 48 3.2.1 General Parameterizations rules ............................................................................................................... 48 3.2.2 Examples of parametric modeling.............................................................................................................. 49 3.2.3 Examples of different configurations ......................................................................................................... 51

3.3 AUTOMATIC MESH MODEL................................................................................................................................. 53 3.3.1 Meshing Tool.............................................................................................................................................. 53 3.3.2 Modification of CATIA model for making automatic mesh possible .......................................................... 57 3.3.3 Parameterization of the meshed product.................................................................................................... 59 3.3.4 The Automatic Mesh Environment ............................................................................................................. 60 3.3.5 Examples of different configurations ......................................................................................................... 61

4 TRANSLATION PROGRAM................................................................................................................................ 62 4.1 FUNDAMENTALS ................................................................................................................................................. 63 4.2 IMPLEMENTATION............................................................................................................................................... 64

4.2.1 Environment ............................................................................................................................................... 64 4.2.2 Results ........................................................................................................................................................ 66 4.2.3 Validation................................................................................................................................................... 72

5 DISCUSSION AND CONCLUSION ..................................................................................................................... 76 5.1 THE THESIS WORK OVERALL ............................................................................................................................... 77 5.2 PARAMETRIC MODEL.......................................................................................................................................... 78 5.3 PANAIR............................................................................................................................................................. 80 5.4 TRANSLATION PROGRAM .................................................................................................................................... 81

5.4.1 Comparison between different PAM configurations .................................................................................. 81 5.4.2 Comparison between different Cessna surface configurations .................................................................. 82 5.4.3 Validation................................................................................................................................................... 83

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6 APPENDIX .............................................................................................................................................................. 86 6.1 CATIA ............................................................................................................................................................... 87

6.1.1 Examples of different mesh configurations ................................................................................................ 87 6.1.2 Fuselage plus Wing configuration for PANAIR analyzes .......................................................................... 92 6.1.3 Full Configuration minus the Engine for PANAIR analyzes ...................................................................... 97 6.1.4 Pre-set Engines ........................................................................................................................................ 104

6.2 TRANSLATION PROGRAM .................................................................................................................................. 105 6.2.1 Edge Numbering ...................................................................................................................................... 105 6.2.2 Forced Intersections................................................................................................................................. 111 6.2.3 Input File of the “Full Configuration minus the Engine”........................................................................ 113 6.2.4 Output File of the “Full Configuration minus the Engine” ..................................................................... 119

6.3 SURFACE DIVISION VALIDATION RESULTS ....................................................................................................... 120 6.3.1 Lift Coefficient results (Cl) ....................................................................................................................... 120 6.3.2 Induced Drag Coefficient Results (CDi) .................................................................................................. 121

7 REFERENCES ...................................................................................................................................................... 122

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Notation CATIA V5 CATIA Computer Aided Three dimensional Interactive Application Curve Smooth Smoothes out a curve from sharp edges Extrusion Gives thickness to closed sketches, splines or lines Law A function which can be used to create splines Multi-Sections Surface A function which makes a surface by the use of two cross-sections and help lines Rule A function in which a programming script can be inputted Reaction A function in which a programming script can be inputted Revolve A function which makes a surface by revolving a line/spline around an

axis Sketch Sketch tool found in CAD programs, usually used for Extrusion

functions. PANAIR 502i $ Marks the beginning of a data block * Identifies the start of a data sub-block = Introduces a comment line Abutment A curve where two or more networks edges meet (exactly or

approximately) alpc Compressibility angle of attack, degrees alpha Angle of attack, pitch (degrees) beta Sideslip angle, yaw (degrees) betc Compressibility angle of sideslip bref Total span of the wing cref Length of the Mean Aerodynamic Cord (MAC) dref Total body length eat Abutment tolerance kn Number of networks input for a specific data block

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kt The boundary conditions to be applied to each network Mach Free stream Mach number Network Rectangular array of panel’s corner points; basic unit for defining the

geometry of the configuration nm Number of points in a network nn Number of point columns in a network sref Full airplane reference area xref X component of the moment reference location yref Y component of the reference location zref Z component of the moment reference location OTHERS Aileron Control surface on the wing which controls the roll BWB Blend Wing Body Aircraft Elevator Control surface on the horizontal stabilizer which controls the pitch Fairing The body that connects the wing and the fuselage Flaps Controls surfaces on the TE of the wings which are used during

landing and take off Leading edge (LE) Front edge of the wing/tail Pitch Rotation around the Y-axis Pro Engineer Computer Aided Design Program Pylon The body that connects the Engine and the wing/fuselage Roll Rotation around the X-axis Rudder Control surface on the vertical tail which controls the yaw Slats Controls surfaces on the LE of the wings which are used during

landing and take off Trailing edge (TE) Back edge of the wing/tail Yaw Angular movement around the Z-axis

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1 Introduction

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1.1 Background Aircraft design is a compromise between many competing factors and constrains. These constrains are mainly economical and technical, both having a great influence on how a design is carried out. The technological depends on the economical, therefore it is necessary to find new methods that will allow engineers to lower the time that takes to develop a new design and at the same time lower the cost. The current challenges in the aeronautical industry are to offer better designs and or methodologies at a lower cost and improve design and production time. The scope of this master thesis is to propose a tool that will have an impact on the early design stages; it will be done by implementing an interface between a CAD model and an aerodynamic analysis program. By doing so the time spent during conceptual and preliminary phases for a new design project should be reduced.

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1.1.1 Conventional Design Process The design process is the means by which the competing factors and constrains which affect the design are synthesized with the specialist analytical inputs to produce the overall configuration [2]. The process may be considered in three different parts:

Conceptual design Preliminary design Detailed design

Fig 1.1[2] shows the design process in a typical schedule diagram, where the various phases are not sequential and tend to overlap each other. For this thesis project the area of interest lies with in the dotted line.

Fig 1.1[2] Conceptual Design During this phase conventional and novel configurations are submitted, every proposal is analyzed. The level of detailed analysis depends on the amount of time that the design team has. The novel proposals may require research studies to quantify the effectiveness of the proposal. In the past, conceptual design methodologies were based mainly on algorithmic computer programs. These programs were usually structured on performance relationships, weight estimation relationships, and empirical/mathematical models for aerodynamics and propulsion evaluations. All these relationships are managed by means of parametrical analysis and or optimization algorithms to reach a level of design useful to point out differences among configurations.

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Fig.1.2 [5]

Fig 1.2 [5] is an example on one of the techniques used by designers that want to estimate the wetted area of a new design comparing it with known configurations, this is a statistical approach, if the proposal that is been worked on is a conventional design then the available statistical information will give a rough approximation. Therefore the necessity to develop improved methodologies, this will enable designers to assess different configurations. Preliminary Design At the end of the conceptual design phase every concept will have been studied to a certain extend. During this stage the concepts that are considered having a high risk factor will be eliminated. The ones that do go through will go under a careful scrutiny to evaluate the best design. One or two designs that are selected will go in to the preliminary design stage. The objective of this second phase is to find the optimum geometry for the aircraft with regard of the commercial prospects and in comparison with competitor aircraft. During this stage the chosen design will be submitted to a more rigorous technical analysis, where the principal parameters are considered to be variable during this analysis. With the advancement

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of computer power, there has been a great deal off improvement in CAD programs and analysis programs. These programs have been integrated into design methodologies and in to this stage as well as in the conceptual phase. Detailed Design When the decision of entering a full scale development, the final detail design phase begins in which the actual pieces to be fabricated are designed, in this way the structure must be defined in complete detail, together with complete systems, including flight deck, control systems, electrical and hydraulic systems, landing gear, cabin layout, etc.

Fig 1.3 [1]

The author of this reference [1] describes the way the aerodynamic design process should be embedded in the overall preliminary design. It also shows how the process in its different stages becomes an iterative process going back to the CAD block for its re-evaluation after analysis.

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1.1.2 Objective of this thesis The main purpose of this thesis work is to lay the foundation of another way to conduct Conceptual Aircraft Design. The objective pursued during this thesis work is to integrate two powerful tools that have been used separately during the design process by some parts of the aircraft industry. This integration of a CAD design program and Panel Code analysis software will then be the first stage of a new methodology. The next step after this thesis work is to implement an optimizing script and make a parametric structure for the empty Aircraft shell.

Fig. 1.4

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Fig. 1.5 Integration of Conceptual design with Preliminary design

The objective of the current project, however, is to develop an interface between the CAD program CATIA V5 and the PANAIR 502i analysis program. The CAD model should be parametric and be able to change in a wide variety of civil jet aircraft. The interface should also have the capability of being mostly automatic. A clear consequence of this methodology should be the integration of both, Conceptual and Preliminary design stages. But having in mind that this integration will not be complete, there will always be some parts that will be characteristic and exclusive of each stage.

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1.1.3 Planning For efficiency purposes the thesis work and research has been divided accordingly:

Mehdi Tarkian Generative Shape Design, Advanced Meshing Tools and Knowledge Advisor Modules research in CATIA. Parametric Aircraft Model Automatic Mesh Model Translation program

Francisco J. Zaldivar T. Mesh Generation tools Panel Code Research PANAIR research and application

It is fair to say that during the development of different sections that comprehend this project the authors have worked together to find the best solution possible to the problems encountered during its development. The results of each section have had a direct impact on the project as a whole.

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2 PANAIR 502i

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2.1 Fundamentals The panel code program that has been used during the development of this chapter is PANAIR. This chapter is focus on understanding what the PANAIR program does, how the model to be analyzed is to be inputted in to the program and what information is obtained from the program results file. At this point is good to acknowledge that because of the extension of the definition and explanation of the input files and output files, only the parameters and results that have been regarded as the more relevant, are the ones that are going to be explained during the duration of this chapter.

2.1.1 What is PANAIR PANAIR (an abbreviation for “panel aerodynamics”) is a robust computer program developed to predict inviscid subsonic and supersonic flows about an arbitrary configuration by means of a higher order panel method. Generally speaking, a panel method solves a numerically linear partial differential equation (the Prandtl-Glauert equation) by approximating the configuration surface by a set of panels on which unknown “singularity strengths” are defined, imposing boundary conditions at a discrete set of points, and thereby generating a system of linear equations relating the unknown singularity strengths. These equations are solved for the singularity strengths, which provide information on the properties of the flow. Higher order panel method implies that the singularity strengths on each panel are not constant. That is why this method is needed specially for solving the supersonic problem. The potential for numerical error is greatly reduced in the PANAIR program by requiring the singularity strength to be continuous. It is also this “higher order” attribute that allows PANAIR to be used to analyze flow about arbitrary configurations. PANAIR can handle simple configurations as well as complex configurations. Most problems can be modeled with a minimum of user input. In general, the aircraft surface is partitioned into several networks of surface grid points (mesh), such as a forward body network, a wing network, and so forth. The coordinates of the input grid points must be computed and entered by the user PANAIR does not generate grid point coordinates. PANAIR connects the grid points in each network with piecewise flat panels. The user also supplies information concerning the free-stream onset flow, the angle of attack, and the sideslip angle. Numerous flow quantities are computed at points on the vehicle surface and at points in space. These include pressure coefficients, values of velocity, forces and moments, to mention a few. The pressure coefficients on individual panels are fitted with two-dimensional quadratic splines and integrated to obtain the six components of force and the moment coefficients. These coefficients may be output for each panel, for columns of panels, for each network, or for any combination of networks. The user has extensive control over the type and quantity of data that is output during a PANAIR analysis.

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2.2 Implementation

2.2.1 Working with PANAIR PANAIR is a panel method program; this means that the geometry that is going to be analyzed needs to be meshed with quadratic elements. An important limitation to consider in representing a surface by network points is that, except for collapsed network edges, all panel sides have a finite length. A collapsed network edge can be consider as a triangular element, these triangular panels or collapse edges are use to start and end, a network such as the nose section of the aircraft.

Fig. 2.1 [3] a) Starting collapsed network b) Starting and ending collapsed network

Fig. 2.2 Surface CAD model of a small business jet configuration

Fig 2.3 Meshed CAD model of a small business jet configuration

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PANAIR is dimensioned to handle model up to the maximum numbers listed below:

Item Maximum Allowed number Networks 150 Panels 18000 All network points 20000 Panels per network 8000 Control points 24000 Total number of points in an abutment 200

Table 2.1 [3] Taking this information in to consideration the model has to be divided into different networks and meshed. Each network will have an assigned name that will help to evaluate the results obtained, as well as track potential errors that may occur. Also it is important to say that understanding how the configuration needs to be divided will help the designer understand how the CAD model should be constructed.

It is very important to follow a logical partition of the geometry; this logical partition will make it easier at the time of imputing the geometry in to PANAIR. A proposed way to partition the model is:

Nose Front mid section Mid Section Tail cone Wing Horizontal stabilizer Vertical tail Pylon (engine mount) Engine housing

This is just one way of dividing the geometry; the actual division will depend mostly on the meshing program that will be used. Fig 2.4 shows an exploded view of a mesh aircraft according to the division presented above.

Fig.2. 4 Exploded view of the different parts forming the CAD model

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2.2.2 Input File The PANAIR program is written in FORTRAN, therefore the input file needs a file extension of .INP, denoting by this that this will be the input file. This file contains several data blocks that have a particular function; they can be divided in to three main parts:

Initial parameters Geometry Section properties

The different data blocks contain the necessary inputs to solve a flow boundary value problem, calculate flow-field properties and control the output information. Here they will be explained briefly, having an emphasis in a few parameters that are important to the reader to understand the methodology that has been followed during the duration of the project. The data-blocks that are not discussed here can be found in the user’s manual reference [3]. INITIAL PARAMETERS – the following data blocks compose this section:

$TITLE $DATACHECK $SYMMETRY $MACH NUMBER $CASES – No. SOLUTIONS $ANGLES OF ATACK $SIDESLIP ANGLE (yaw) $REFERENCE FOR ACCUMULATED FORCES AND MOMENTS $PRINT OUT OPTION $BOUNDARY LAYER $VELOCITY CORRECTION.

At this stage the user defines the primary analysis parameters such as, angle of attack, yaw angle, velocity (Mach number). Other extremely important parameters during this stage, that defines the output file and basic information of the aircraft to be analyzed. These parameters are taken form the aircraft it self:

Full wing reference area (sref) Span (bref) Mean Aerodynamic Cord (cref) Body length (dref) X Y and Z components of the moment reference location (xref, yref, zref, aircrafts

aerodynamic center) The configuration forces and moments summary gives the lift coefficient (Cl) induced drag coefficient (CDi) side force (CY) and forces and moments about the reference coordinate system (FX, FY, FZ, MX, MY, MZ) for both the inputs and the complete configuration. There for, it is crucial that the designer has the information about where the aircrafts aerodynamic center is located in order to obtain the results with a higher degree of accuracy from the program.

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The boundary layer and the velocity correction control input data blocks initiate velocity corrections for generating boundary layer input data and fro internal flow (for example, an inlet) these corrections primarily improve the flow properties for flow slower than the onset flow. The biggest corrections are made near the stagnation region. The correction changes the surface velocity components, mass flux components, pressures and Mach numbers. The print out option is of great importance, because they will give the necessary information to the designers about the configuration that they are studying. They will be explained with more detail further in to this chapter. To see an example of the first block please refer to the appendix 6.2.3 GEOMETRY – the following data blocks compose this section:

$POINTS $TRAILING WAKES $PEA – PARTIAL OR FULL EDGE ABUTMENTS $EAT

This is the data block where most of the problems occur: a full model configuration may contain more than 7000 grid points, making a time consuming task and prone to error. In order to get to the arrangement of the grid points of every network it is necessary to analyze the network that are been introduce to the program, that is the first thing, the second thing to do, is to observe how the meshed network have been divided, with this information it is possible to determine the number of rows in the network (nm, see the notation section in this chapter) and the number of point columns in the network (nn, see the notation section in this chapter). The values of nm and nn will depend on two things, the first would be the size of the mesh and the second would be the direction which the designer decides to use, meaning, where the network will start and in which direction.

Fig 2.5 Grid point’s extraction from meshed surface

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Fig. 2.5 is intended to explain in a more graphical way in which the grid points of a network are to be organized; the red numbers represent the panels order, and the numbers in black are the grid points of the panels. Worth mentioning is the fact that when you have a meshed surface with 9 X 8 panels, then the nn x nm matrix will always be Rows+1 x Columns+1. This characteristic of the grid points will allow us to visualize in a clearer way the network construction. The grid point matrix that will go in to PANAIR to represent the geometry will have the following format: x(1,1) y(1,1) z(1,2) x(1,2) y(1,2) z(1,2) x(1,m) y(1,m) z(1,m) x(2,1) y(2,1) z(2,2) x(2,2) y(2,2) z(2,2) x(2,m) y(2,m) z(2,m) …. …. …. X(n,1) y(n,1) z(n,2) x(n,2) y(n,2) z(n,2) x(n,m) y(n,m) z(n,m) To analyze the flow about a configuration, you must describe geometrically the surface boundaries and specify boundary conditions that best represent the physical flow. PANAIR is capable to use a great variety of boundary conditions. The boundary conditions that fulfill the needs for the analysis of the models that concern the scope of this work are:

Kt= 1 Indirect condition on an impermeable thick surface; preferred for satisfying impermeability, wings and body

Kt= 5 Base surface condition; used to represent the aft blunt base on a body or wing with a thick trailing edge.

Kt= 9 Flow through surface, fan face, inlet, etc; commonly used to represent flow into or out from a surface.

Kt= 18 Vorticity matching Kutta condition used for sharp trailing edges; commonly used for wings with different upper and lower surfaces.

Kt=19 Calculates higher-order wakes.

Kt =20 Constant strength doublet wake; commonly used as a connector or filler wake between the wing wake and body wake.

After the configuration has been defined the next step is to declare the wakes that follow the geometry. Wakes are attached to a network edge and go straight downstream (y = constant and z = constant) to the specified x wake coordinate. All input wakes should be terminated at some common, convenient location aft the configuration. The data block to define the wakes is $TRAILING WAKES, but it is not the only way in which a wake can be inputted in to the program. You can use the $POINT – data block, in this case the coordinates of the wake are given by the designer, this wakes need to start at the trailing edge and go down stream for each column of the network.

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a)

b)

Fig.2.6 Wakes forming behind an airplane; a) top view; b) side view

In the lines above there is mention of some very important concepts that need a good understanding, because the concept of and edge in PANAIR is a subject that, if not properly considered, will generate errors that could stop the analysis program. The edge of a network is something very important and very simple; every network has 4 edges, what is important to know is how to identify each one of these edges. In the following images there are 2 examples on how to recognize the edge number of a network. The first image is a rectangular network, which is the most common network and the second one is a triangular network, this one can be found on the nose section of a fuselage. The importance on having a good understanding on the network edge definition lies with the abutments, since in order to abut the networks it is necessary to define the edges that are side by side and which points are to coincide with each other.

Fig. 2.7 [3] rectangular network arrangement

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Fig. 2.8 Triangular network arrangement (collapsed network) After configuration geometry has been described and the proper boundary conditions have been defined, the next important task is to check the networks abutments. The abutments provide the means for PANAIR to maintain the continuity across network edges. When there are two separated networks that coincide at their edges, it is necessary to verify that there are no gaps between these two networks, but, because the meshing has been done with planar quadratic elements, non matching panels are bound to occur. It is very difficult to avoid them, but never the less, it is the responsibility of the designer to review and accept the abutments as determined by the program, and keep the non matching panel corners to a minimum. As the heart of the program requires, the abutting network edges must have exact panel edge points which match along the networks edge or panel edge points which are on the straight line between the exact points. Thus, the interfaces of two or more doublet network surfaces must match, i.e., have no gaps between adjacent networks. The user must keep non-matching points to a minimum with only small mismatches, and use the network option to retain the original unit normal vector (cpnorm = 2.0) Basic assumptions concerning abutments are listed below

All network edges abutments are assumed to start and end at identical network edge points.

To maintain the quality of the original input geometry along an abutment, match as many points as possible.

The user is responsible for reviewing and validating abutments before making a solution run.

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Fig. 2.9 Abutments between the wing and the fairing networks

SECTION PROPERTIES – the following data blocks compose this section:

$SECTIONAL PROPERTIES $FLOW PROPERTIES $XYZ COORDINATES OF OFF-BODY POINTS $GRID – OFFBODY POINTS $STREAMLINES $TREFFTZ PLANE ANALYSIS $END

PANAIR can compute mass flux or velocity components, second order pressure coefficient and perturbation potential for:

Points in flow field input as: Points Grid points

Streamlines traced in flow field In this field the program has limitations accordingly to the number of solutions that the user has asked for (maximum 4 solutions); for one solution less than 1666 points; two solutions 1000 points; for three solutions 666 points; four solutions 500 points; and the total number of streamlines must be les than 500. Another feature from this data block is that the user can define a number of plane cuts, this cuts allow the designer to observe the behavior of parameters such as pressure, velocity, forces, etc. at

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specific regions of the model or network. For example if the designer needs to pay attention on the wing tip, or at the MAC, etc.

Fig. 2.9 a) Isometric view of an arbitrary cut; b) top view of a wing cut with constant y [ ]

The program has the capability to analyze the surrounding flow around the aircraft, it is interesting to observe this changes depending on the configuration that it is been studied; the behavior of the flow around the geometry will enable the designer to have a greater understanding of the design he or she is working on, and more importantly, having the option of analyzing the surrounding flow will enhance the results obtained by the computer. To analyze the surrounding of the geometry the program uses two methods, off-body points and streamlines. The off-body points are a cloud of points that surround the body, depending on how dense the designer wishes to make this cloud and how many points he inputs, as it has been said in the lines above there is a limit on the number of off points that can be used.

Fig.2.10 Streamline representation

And a streamline is a path traced out by a mass-less particle as it moves with the flow. Fig. 2.10 shows the computed streamlines around an airfoil. Since the streamline is traced out by a moving particle, at every point along the path the velocity is tangent to the path. If the velocity and the pressure at different points of the streamline are known, then, the behavior of the flow around a body can be visualized.

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This information will result valuable for the aerodynamic designer at the end of the day, because by visualizing the streamlines over the body it will be possible to see if the flow is possible or not. That is why it is very interesting to study the streamlines around a specific area of the model. The last command in the input file refers to the Trefftz plane analysis [3] [9]. This analysis improves the accuracy of the induced drag calculation as compared with the integrated pressure induced drag, but on the other hand when referred to lift calculation compared to surface pressure integration, the second method is found to be more accurate. After reviewing the more significant data blocks of the input file it is fair to state some recommendations. When working with PANAIR, the most time consuming task for the PANAIR user is the meshing of the model. To obtain a suitable paneled geometry some guidelines should be followed:

The panel distribution needs to be considered in such a way that you do not end up having a large panel next to a small panel, for a better result they need to be gradually larger or smaller.

To obtain a better pressure distribution analysis, create a denser mesh where you know the pressure is greater, and you can have a more relaxed mesh where the pressure distribution is not that important.

Always try to avoid un-matching panel corners; it is important to reduce the number of forced abutments along the networks. Avoiding gapes between connecting networks will improve the quality of the result.

When defining a network it is very important to avoid having a change in direction grater than 20 degrees (Fig. 2.11b))

Before introducing the model in to PANAIR the connecting networks should be check to see if there is need for forcing abutments, if so see the possibility to modify the network in order to simplify the input file.

The panels that are to be created with the mesh of the model need to have an aspect ratio grater than 0.001. The panel aspect ratio is the length of the panel divided by its width

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Fig.2.11 Examples for suitable model paneling a) Correct and incorrect way to distribute panel density b) Direction change in a wing network.

2.2.3 Output File The printout file represents the sequential processing of the program. A large portion of the initial printout provides information on how the program interprets and processes a particular case. The last section of the printout gives analysis results for the different solutions asked in the input file. Of prime interest to the user, are the resultant forces and moments and the detailed pressures at each control point. The file can be regarded as have two main sections; “Data Check” and “Solution Run” and these two sections can be divided in to several sub-sections:

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DATACHECK RUN:

1. List off the input file; Record of the input processing; Quick summary of the input parameters; record of the forced and partial edge abutments.

2. Liberalized geometry, abutment analysis. 3. Geometry before liberalized geometry. 4. Summary of extra control points. 5. Abutment summary. 6. List of network edge points moved by $EAT. 7. Panel data. 8. Control points for network #. 9. Boundary condition information for network #.

SOLUTION RUN:

1. Problem and network indices. 2. Simultaneous solution number #. 3. Force/Moment data for network #. 4. Sectional properties, cut definitions and reference data, solution #. 5. Sectional properties, cut force and moment data, solution #. 6. Sectional properties, network force and moment data, solution #. 7. Input configuration forces and moments summary. 8. Full configuration forces and moments summary. 9. Off-Body flow characteristics. 10. Job cost summary by function.

Knowing the different parts of the print out will make the task of searching and extracting information much simpler. The information that the user needs to have at hand will depend on what he wants to analyze. This is section as it can be appreciated by the list above is extensive, that is why the reader should refer to the manual to read about all the different parts of the output file. Never the less, in the following pages the reader can find a few of the subsections in order to understand the information that is given on such sections and also there will be screen shots of the output file section that it is been refereed to. Since it has been said earlier that the input file is prone to error due to the large amount of geometry information that it is been handled, the error notifications that will stop the analysis run will always be shown at the end of the file, so it is only fair to say that, a good practice when starting to read the output file is to go directly go to the end of the file and check if there are no errors in the run. If there is an error, normally it will be briefly explained. The most common error that has been found to occur will happen when inputting the geometry, simply because, as stated before, the geometry represents a very large amount of data which is been handled. PANAIR is written in FROTRAN, this language is very unforgiving for format errors. When such errors are detected by the program, it will stop the run and the output file will present at the end of the file, a statement of error, fig 2.12.

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Fig. 2.12 Example of error due to ill formatted data

As it can be seen in the figure above, this type of error is very well explained, the error message is giving the type of error (ill-formatted data) and in which line of the input file the error is (line 55). With this information one can go to the beginning of the input file and read the exact line that presents the problem. Also worth pointing out, is the fact that the solver shows where it has stopped the analysis, in this example, it stopped in the “$poi” Data Card, this information also will help locate the source of the error. Another common mistake occurs when defining the abutments between the neighboring networks, this error is harder to locate since the solver starts its calculations and by the time that it has encountered the problem, the output file is already of considerable extension.

Fig. 2.13Example of error due to improper abutment definition

Fig.2.13 shows an error that occurred during the abutment identification, here the program at the end of the page only states that there is an error in abutment identification (abmt = abutment) and there is no reference to any line in the output file in particular it is then the task of the user to find this. The user will then scroll trough the file until he locates the part of the print out that specifies which abutments are not correctly aligned, the information that the program reads is as it is shown in Fig. 2.13. After the program has checked that the data provided is correct, the output file prints, in accordance with the print options given during the first block of input parameters (section 2.3.1, Initial parameters), the results of the analysis divided in the number of cases the user has specified, remember that the number of cases referred to which angles of attack the analysis has

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been defined for. Each case will be given as a simultaneous solution number #, (solution 1 for alpha 1).

Fig. 2.14

Fig. 2.14 shows the start of the solution printout, this header is common to every solution run. The full configuration forces and moments summary is probably the most important section of the output file since it gives the lift coefficient, induced drag coefficient and side forces for all the solution runs that the user has chosen at the beginning of the input file (max number of solutions per run 4). The knowledge of these quantities is of great importance since it allows the designers to compare different proposals during the conceptual design phase, these comparisons then lead to a better decision to which of the designs that have been compared should advance to the preliminary design phase.

Fig. 2.15 Results summary

Fig.2.15 shows how the forces and moment summary looks like, as can be seen this summary shows both solution for Cl and CDi, Surface pressure distribution and Trefftz plane analysis, also gives the total area of the model that is been analyzed.

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2.3 PANAIR Surface Validation

By this time the authors already know what the program needs in order to run the analysis, but actually they did not know exactly how to input the geometry taken from a CAD file. During this section it will be explained what procedure was used to validate the surface division and at the end of the chapter it will be clear which is the proper way to divide the CAD model. Also during this section it will be explained how to extract the coordinates of the grid points that comprise each surface.

2.3.1 Validation procedure To validate the surface division it is necessary to fallow a series of steps to guarantee a credible result. The first step is to have something to compare against that will show how close or fare the results that are been drawn are, the geometry that is going to be used as the basic geometry is Example number 9 from the PANAIR Users Manual [3], it is a simple wing body model (SWB) Fig 2.16

Fig. 2.16 Simple Wing Body model

Secondly, the SWB geometry will be divided in several different ways, which represent how the final CAD model could be divided depending on the capabilities of the meshing tool used to extract the grid points from the model’s surface. A third part for the surface validation will be reproducing the exact same geometry, but this time using a CAD generated model. This CAD generated model needs to be exactly the same geometry in order to have a real comparison between the two sets of tests. The CAD generated model will be identified whit a letter “C” in front of the designated test, for example, comparing the original geometry “SWB” then the geometry generated with Cad would be “C-SWB”. The very last section of the validation procedure is comparing the results that where obtained to be able to give a valid surface division, that will be implemented in the final model. The values that will be compared are:

Lift Coefficient (Cl) Induced Drag Coefficient (CDi)

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According to the users manual, the Cl is calculated using a surface pressure distribution and for the CDi a Trefftz plane analysis method. To show how the values for each coefficient changes the result table will show both coefficients using both calculation method. Before describing how the grid points are extracted from the surface Table 2.1 will describe the different ways (11 in total) that are going to be tested, this including the original model.

Model Description SWB Simple Wing Body, original example [3, chapter 9]

SWB1 2 wing surfaces, inner wing and outer wing

SWB2 2 wing surfaces, upper wing surf and lower wing surf, TE-LE and LE-TE

SWB3 2 wing surfaces, upper wing surf (TE – LE) and lower wing surf (TE – LE)

SWB4 Nose, upper mid body, lower mid body

SWB5 Multi Section fuselage, top to bottom point direction

SWB6 Multi section fuselage (2 aft surfaces), top to bottom point direction

SWB7 Multi section fuselage, left to right point direction

SWB8 Multi section fuselage (2 aft surfaces), left to right point direction,

SWB9 Multi section fuselage (2 wing surfaces), left to right point direction (SWB2+SWB8)

SWB10 Multi section fuselage (2 wing surfaces), top to bottom point direction (SWB2+SWB6)

Table 2.1 Surface division for original model The reader must remember that, besides inputting the surface, the wakes must be defined too, Fig. 2.17 shows the SWB model with the surface division that the user’s manual shows, the panels with the gray color are the wakes.

Fig. 2.17 SWB surface division and wake definition

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To have a valid comparison, the same boundary conditions must apply to all the different configurations, these boundary conditions for the first model are:

Surface name nm nn Boundary condition 1. Wing 11 3 kt = 1 2. Wing Tip 6 2 kt = 1 3. Lower Body 11 3 kt = 1 4. Upper body 11 3 kt = 1 5. Body Base 5 2 kt = 5 6. Body to Wing Wake 4 2 kt = 20 7. Wing Wake - - kt = 18 8. Lower Body Base Wake - - kt = 18 9. Upper Body Base Wake - - kt = 18

Table.2.2 SWB surface brake up for PANAIR analysis

2.3.2 Grid Point Extraction PANAIR is a sensitive program when it comes to the grid point’s arrangement, there for it becomes an important topic to discus in this report. In section 2.3.1 it was mentioned that during the validation of the surface division the SWB model was going to be reproduce using a CAD program, the idea behind this is to ensure that the points given bye such program will render the same results as the ones given by the different configurations in which the original test model (SWB) has been divided. There for a CAD model, which is exactly the same as the original model, was made and divided in the same fashion. Table 2.3 shows the description of the different configurations. Note that they are the same as for the original model.

Fig. 2.18 Exploded view of model C-SWB

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To extract the points from the CAD model, first it is necessary to understand how the points are given so they can be extracted and rearranged. Fig. 2.19 shows a portion of the text file that CATIA (CAD program) produces. This portion shows the type of element for the panel, the label of the panel (panel number), to which surface it belongs (order in which the model was meshed) and the points number that form the corners of the panel.

Fig.2. 19 panel/corner points for a CAD generated model

The points number correspond to the number of point which CATIA has assigned to it and this point is given in Cartesian coordinate form (X Y Z) as it is shown in Fig. 2.20

Fig. 2.20 Point number and coordinates layout in text file from CATIA

Corners on each panel depend on how the mesh is defined (this will be explained further in chapter 3).

Fig.2.21 Panel corner numbering

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Knowing the corner orientation and how the panels are label will make the extraction of the points possible, and will let the designer decide what direction to follow a long the columns in the network (surface). A panel/corner matrix has been done to extract the coordinates for each point on a row. The number of panels that constitute a surface determine how many columns and points in each column the network will have. To illustrate this concept lets take fig 2.22 that represents the nose surface of an aircraft.

Fig. 2.22 Panel numbering for the nose surface mesh

This surface is divided in to 16 panels starting from the point on the far right of the image and going to the left there are 4 panels, and going from bottom to top there are 4 panels as well (4 x 4). This distribution of the panels will give as a result a Panel/Corner matrix of 5 rows or network point column and 5 point columns (mn x nn).

Fig.2.23 Panel-Corner Coordinate matrix

From this matrix (Fig. 2.23) we can observe the pattern that prevails during the extraction of the body points. Thus, the logic in which the points are extracted is the same for every single surface on the model, it depends on two basic things, first the manner in which the mesh is produce and secondly on the direction of the points in the first column. After extracting the points from the panel corner matrix it is mandatory to rearrange the points following the format that PANAIR works with, this format is explained in section 2.2.2 of this same chapter under GEOMETRY.

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2.3.3 Result comparison After running all the different configurations it is time to show the results that PANAIR has given. From these results, as mentioned before, it is going to be decided on how the CAD model needs to be divided in order to get a god result. The analysis that has been run is for the following characteristics:

Mach = 0.6 Compressibility angle of 4 degrees Alpha 1 = 2 Alpha 2 = 4 Alpha 3 = 8 Symmetric in XZ plane Boundary conditions as shown in table 2.2 Surface pressure distribution analysis Treffzt plane analysis

First table will show the results for lift coefficient Cl and the second table shows the results for induced drag coefficient CDi.

ALPHA Surface Pressure Treffzt Surface Pressure Treffzt 2 Cl diff (%) Cl diff (%) Cl diff (%) Cl diff (%)

SWB 0.08159 0.08214 C-SWB 0.08993 0.09335 SWB4 0.08177 -0.22 0.0821 0 C-SWB4 0.09026 -0.37 0.0933 0.01

SWB10 0.08109 0.61 0.08226 -0.15 C-SWB10 0.08085 10.1 0.08259 11.53 4

SWB 0.1628 0.1639 C-SWB 0.17155 0.1759 SWB4 0.16311 -0.19 0.1639 0 C-SWB4 0.17206 -0.3 0.1759 0

SWB10 0.16225 0.34 0.16436 -0.27 C-SWB10 0.16226 5.42 0.16496 6.21 8

SWB 0.32224 0.3267 C-SWB 0.33205 0.3401 SWB4 0.32276 -0.16 0.3267 0 C-SWB4 0.33283 -0.23 0.3402 0

SWB10 0.32228 -0.01 3.28E-01 -0.33 C-SWB10 0.323 2.73 0.3289 3.3 Table 2.3 Cl result comparison table

ALPHA Surface Pressure Treffzt Surface Pressure Treffzt

2 CDi diff (%) CDi diff (%) CDi diff (%) CDi diff (%) SWB 0.0008 0.0013 C-SWB 0.00198 0.00165 SWB4 0.00089 -11.25 0.0013 0 C-SWB4 0.0021 -6.06 0.0016 0.02

SWB10 0.002 -150 0.0013 -0.25 C-SWB10 0.00274 -38.38 0.0013 21.64 4

SWB 0.00682 0.0051 C-SWB 0.0079 0.0058 SWB4 0.00692 -1.47 0.0051 0 C-SWB4 0.00805 -1.9 0.0058 0.01

SWB10 0.00817 -19.79 0.0051 -0.49 C-SWB10 0.00868 -9.87 0.0051 11.63 8

SWB 0.03055 0.0203 C-SWB 0.03121 0.0218 SWB4 0.03069 -0.46 0.0203 0 C-SWB4 0.03143 -0.7 0.0218 0

SWB10 0.03223 -5.5 0.0204 -0.61 C-SWB10 0.03202 -2.6 0.0204 6.15 Table 2.4 CDi result comparison table

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The difference column is given in percentage and it is compared to the results from the original model SWB for the pressure surface integration method and for the Treffzt plane analysis method respectively. This will be true for the CDi comparison table as well.

The configuration SWB5 and SWB7 shown in appendix 6.3.1 and 6.3.2 use a different boundary condition for the body wing wake (kt = 19) this condition is for higher order trailing wakes, it was necessary to change the boundary condition in order to get a result from the solver. The reason why the normal boundary condition (kt = 20) did not work is because the trailing edge was not attached to a network edge, this was because the aft section was configured in too one single surface. And in the other configurations the aft section was divided in too two section providing a network edge to attach the trailing wake. Also it is good to point out for the reader that, using a boundary condition of kt = 19 in the remaining configurations did not change the result that we observe in the previous tables. After making this comparison tables for the two coefficients the conclusion that can be drawn are that there are two possible configurations that can be used, these two configurations are shown in the following table: These two configurations are possible since they are the closest in percentage to the original result. When the final CAD model is complete and the meshing tool is fully understood then it will be possible to say which configuration will be used. In the mean while the following figure shows the two configurations to be considered.

Fig. 2.24 C-SWB4

Fig. 2.25 C-SWB10

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3 CATIA V5

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3.1 Parametric Aircraft Model

3.1.1 About CATIA in this thesis As a student of Linköping University one is well familiar with the CAD program Pro-Engineer. There are several fundamental courses and also courses which are more in dept. The author took many of these courses and was quite satisfied with the results of this program when working with Part and Assembly design. However working with surfaces was a frustrating task. To even be able to create a surface one has to create the lines necessary aligned in a very specific way, furthermore once a surface was made, it was nearly impossible to modify it even slightly. The author was introduced to CATIA in the last 6 month of his education and realized that creating and modifying surfaces in this CAD program was much easier. Therefore the choice between the two was quite simple, even though as mentioned above, his experience in Pro-Engineer exceeded CATIA several folds. The modules chosen to work with during this thesis work have been

1. Generative Shape Design 2. Advanced Meshing Tools 3. Knowledge Advisor 4. Assembly Design

The author has chosen to split up the different sections of the Aircraft Model into parts such as “Cockpit”, “Front fuselage” etc. This has been done for this specific model and in no way represents the real terms which occur in the aeronautic industry. Furthermore it should be mentioned that the author only refers to half of the aircraft (with symmetry on the ZX axis), so terms such as “one engine under wing” should be interpreted as two engines for the entire aircraft model.

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3.1.2 Building structure One main reason for building the Parametric Aircraft Model was to being able to export the geometry by the means of mesh generating, to PANAIR. However the model has shown to have the potentials to carry out structural analysis as well. Trying to accomplish these goals, the model was built and re-built several times. Mostly the main reason has been that it hasn’t been simple enough in a design and parametric point of view. Complex solutions usually give an error somewhere along the way. Therefore besides not using some tools at all because they are bound to give a failed parametric model, one of the aims has throughout the thesis been to keep it as simple as possible. This cannot be emphasized enough though it has shown to be one of the main keys to a fluently working model. The model is divided in several parts as seen in Fig 3.1.

Fig 3.1

Each part has its own set of parameters which are controlled by a set of customized Rules and Reaction. As a parametric model, it is only needed by the user to change the needed parameters to get a desired result. The external surfaces of an airplane not included are in this thesis work are winglets, slats and flaps among other things.

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3.1.3 Fuselage Part

Fig 3.2

When opening the fuselage part one will see a list of parameters (as seen above), which are made to control specific geometries. Such as the “Radius” parameter for example which, as the name suggests, controls the radius of the fuselage. All parameters are directly or indirectly under supervision of a set of Rules and Reactions which makes sure that the user stays in the frames of the CATIA rules and that the model, ultimately keeps the shape of an Aircraft model.

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The parameters “Front Fuselage”, “Middle Fuselage” and “Tail Cone” control the length of the cabin and the end of the fuselage. As can be seen in the picture 3.2, they have the values of 0.25, 0.35 and 0.35 respectively, which combined amounts to 95%.

Fig 3.3

If the user were to change one of the parameters above by more than 5% units which would result to a length greater than the one being inputted as “Aircraft Length”, one of the made Reactions in this case would warn the user, as can be seen in Fig 3.3. The surfaces of the Fuselage part have been built mostly with Multi Section Surface, Blend and Revolve functions.

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3.1.4 Wing Parts

Fig 3.4

The wing is built in three different parts, being the Inner-, Middle- and Tip- wing as can be seen in Fig 3.4. These parts have more or less the same parameter structure. All of them have a NACA4 (see Fig 3.5) profile, but could easily be modified to NACA5 or any other profile, with the use of the right mathematical formula describing the airfoil in question.

Fig 3.5

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The NACA4 formula is added under a Law, which is automatically modified when changing the parameters given under each part. Some of the main differences between the parts are that for example in the Inner Wing Part one can decide to have the low or high wing configuration and in the Middle Wing Part you can set the dihedral angle of the entire wing. Apart from small differences in the parametric sections, the wings are very alike, CAD construction wise. Please note that the methodology of the wing construction is followed by earlier research conducted in Linköping University.

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3.1.5 Horizontal Tail

Fig 3.6

The Horizontal Tail also has a NACA4 profile. What differs between the build up of the Horizontal tail and the Wing parts is that it has a parameter which allows the user to choose between a conventional Tail and T-tail. Another unique parameter for this part is one that gives the tail a desired deflection as an elevator, typed in by the user of course, under the parameter “Deflection” (see Fig 3.6). It is made in such a way that a point is created on 40% of the root chord from the leading edge. The Horizontal Tail is then made to rotate around this point. This point is not parameterized, though it hasn’t seemed to be a necessity for the user to modify the point. However if required this point can be parameterized so that the user can for example change it to 50% of the root chord.

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3.1.6 Vertical Tail

Fig 3.7

The Vertical Tail is limited to only a symmetric cross-section (, unlike the other airfoil parts,) with a variable thickness. Other than the above mentioned, the Tail part doesn’t have any other distinguished characteristic parameters compared to the wing parts, besides one parameter called “Tail tangency” (see Fig 3.7) which blends the tail into the fuselage and can be modified to have a more or less blended characteristic.

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3.1.7 Engine Configuration

Fig 3.8

In the Aircraft Product tree there are a set of parameters which decides the number and type of engines as seen in above picture. The user can choose between 1-3 engines under the wing, an engine in the back of the fuselage and a tail mounted engine. In the Fig 3.8 the choice is set on 1 engine under the wing.

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3.1.8 Engines choice and positioning

Fig 3.9

Each engine has its own part and in every part you can choose between a wide range of pre-set engines like the Pratt & Whitney PW-4000-94, General Electric CF6-80C2 and Honeywell TFE731 etc. A complete list can be found in Appendix 6.1.4.

Fig 3.10

The engines housings are much simplified to make the model light and easy modifiable. The front consists of a flat plate (instead of blades) and so does the rear of the engine. The rest of the engine is made with the function revolve. When the parameter Engine Choice is changed (see fig 3.9) a Reaction will change the engine configuration and notify the user of the change as seen in Fig 3.10.

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The pylon was one of the absolutely hardest parts to make on the whole aircraft, though it had to be more or less blended in the wing and engine, but at the same time be very modifiable so that when changing engine from a larger engine to a smaller one, the pylon should not go into surface failure. Following picture shows how stretchable the pylons can be, demonstrating an engine change from the Rolls-Royce Trent 800 to a Williams International FJ44-4A.

Fig 3.11

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3.2 Parameterization

3.2.1 General Parameterizations rules Parameterization is more about than just creating a line and letting a parameter determining its length. One must always take into consideration that it is ultimately surfaces we are dealing with and surfaces are doomed to go into failure and finally error the whole product. Therefore special care must be taken. When one parameter is changed then other parameters may be needed to be changed as well and here we are talking about having links all over the whole aircraft model that if used correctly can give satisfying results, but if not approached rightly, it will be a nightmare to operate. The key is to know the limitation of each surface and make sure they never reach those limitations. To achieve this, the Rule and Reaction functions in CATIA are used, which are mildly explained in previous section and more so in the following one. One cannot be protected against every possible error. The user can work in very unorthodox ways, for example wanting to create a vertical tail with a length of 100m (possibly by mistake). It is of course possible to make a Rule against such a scenario, but these exaggerations are endless and going with the methods used in this thesis work, one cannot be protected against all of them. However with CATIA being such an endlessly huge program, it is hard to say if other functions exists which are more global and can handle these kinds of problems better. Furthermore, one should always try to avoid using following function on CATIA V5 when designing a parametric model, involving surfaces:

Parallel line, if not used with a law Try to avoid the sketch function totally. However some functions are worse than others in

sketch such as rounded arc and corner functions. Curve Smooth function is a very tricky one. Special care is required using this function.

So try to avoid if possible. There is a 3D module called FreeStyle. Avoid this module entirely, though it is unfit for

parametric design.

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3.2.2 Examples of parametric modeling The engines consist of several splines which are revolved around an axis. This axis is then controlled by the parameters which decide the position in the x, y and z axis. The user is able to change the position of the engines as long as the inputs are reasonable. However the Engines which are under the wing (Fig 3.11) have two different positioning points, one which is on the wing and has the parameter which moves the engine along the wingspan (*) and the other which is the axis (**) of the engine which controls the x and z position in relation to the first point (*).

Fig 3.11

As mentioned in the previous section, a range of engines (housings) are available under the parameter “Engine Choice”.

(**)

(*)

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Fig 3.12

For the conventional tail there are parameters which control the position of the Tail (Fig 3.12), unless they are out of boundary of the Rules and Reactions, which will change the input into a more suitable one automatically. If the user still wishes to use that specific input, he/she must first change other parameters which will/may make the change possible. One exaggerated example is if the user wishes to have a root chord of 6m for the Horizontal Tail when the fuselage is 10m long. This is and will remain unacceptable for the set of Rules and Reactions in this part, unless the user changes the length of the fuselage (extensively so regarding this example). If the user chooses to have the T-tail configuration, a Rule will make the parameter which controls the position in x direction inactive and the user will get a static horizontal tail. The reason for why the author has chosen to make it static has to do with the later mesh processing.

x direction

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3.2.3 Examples of different configurations The Pictures from this and the following page shows a variety of different configuration and shapes the Parametric Aircraft Model can take. This product has been made in such a way which makes it possible for the user to develop any desired aircraft as long as it is in the frames of the conventional civil/business jet design, meaning that this model can not generate unconventional designs such as a BWB concept or a concept similar to the Sonic Cruiser. The reason for why these designs weren’t implemented in the Parametric Aircraft Model is not because of there complexity but the lack of time to implement such a wide range of different configurations. However it should be underlined that it is by all means not an impossible task to have a much dense range of configurations in one and single Parametric Model. [For the source of the pictures to the left see Reference 16]

Fig 3.13Cessna Citation

Fig 3.14 Airbus 310

Fig 3.15 Airbus 340

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Fig 3.16 Boeing C-17

Fig 3.17 MD-11

Fig 3.18 Boeing 777

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3.3 Automatic Mesh Model

3.3.1 Meshing Tool Earlier it has been said by the authors that one of the main differences between Panel Code Methods and Finite Element Fluid Dynamics is the mesh generation, which is a very important and time consuming task. The advantage of a Panel Code over CFD lies on the mesh, Panel Code as said before, needs only to mesh the geometry that is going to be analyzed (the aircraft) and CFD meshes the volume surrounding the geometry to be analyzed. This brings us to this chapter, where the mesh generation and the tools to do it are to be explained. The panel code mesh works with a structured mesh using quadratic elements, and for very specific sections, triangular elements. For PANAIR, triangular elements are considered collapsed quadratic elements (Fig. 2.1). To do a structured mesh over the geometry involves knowledge of how the panel code program asks for the geometry to be inputted (revised in chapter two), time, and knowledge about the program that it is going to be used to produce the meshed part. One can use a variety of software’s to produce this structured mesh, an example of these programs is Gridgen by Pointwise, this program is reliable source for this task, it is compatible with CATIA V5, ProEngineer, just to mention a few. It is important to mention the different programs that can be used for the task at hand in order to have an informed decision and to be able to evaluate the options that the user could encounter Fig. 3.19.

Fig. 3.19 Mesh generation software comparison table [17]

This information helps to find the best mesh generation program for the project that the user is working on. Other things to consider during the evaluation are the ability to work with CAD files such as CATIA. The task of importing the geometry from a CAD file is of major importance,

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some times when doing such task there is information of the geometry that is lost, there for, having a meshing option in the CAD software would make the meshing section of the project a simpler assignment, but this only motivation is not sufficient to use the built in meshing tool. The meshing tool needs to be capable of performing the mesh with the proper elements, according to the analysis program that is going to be used. Also, the meshing tool has to be capable of working with surfaces and export the grid points of the meshed surface to a text file or spreadsheet file so they can be used in the proper analysis software. Using the static analysis module from CATIA V5 has been found to be, at the moment, the best option to perform the task at hand. Basically, the advantage is that the module already exists with in the CAD program, and there for, there is no need to export the model to the meshing software, ergo, no lost of information on the geometry. And then after researching the programs and evaluating the Advance Meshing Tools in CATIA V5, the authors decided to go forward using the meshing tool from CATIA V5. This tool works with surfaces, it is simple to use, it is advance enough to work with the geometry at hand, and the most important feature that makes it the best candidate at the present moment is the fact that it is with in the CAD program and the user does not need to export the geometry and there will be no lost information in the process.

Fig.3.20

The Advance meshing tool supports quadratic, triangular and unstructured elements. Also it has the flexibility to modify the mesh density, a feature that proved useful, just to remember during chapter two it was recommended to space the panels in a manner in which it was avoided to have large panels next to short panels as much as possible. Therefore, the flexibility of the program that this feature provides suited the project and reinforced the decision to use this tool.

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Fig.3.21 Mesh comparison

Fig. 3.21 shows an example on how the same part can have to different spacing’s of the elements, keeping the same number of elements. There are several functions to edit the number of elements that compose the mesh, and how they are spaced between each other. These options are:

Uniform Arithmetic Geometric User Law

An important feature of meshing the surface of the model that needs to be analyzed is the arrangement of the panels; the user needs to have a full knowledge on how the panels are arranged along the surface. This arrangement actually is decided at the time that the mesh is constructed in CATIA. This starts to take shape at the moment of specifying the global parameters of the mesh; fig 3.22 shows a red arrow, this arrow is defined depending on the construction of the surface and will indicate how the panels will move along the surface (counter clock wise or clock wise).

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Fig. 3.22

After this step is done, the user needs to select the type of element to use, in this case he will select a mapped mesh (quadratic element) of if the surface that is been worked with is the nose of the aircraft he will use a mapped mesh for triangular domain. What the program refers to as “mapped” is for a structured mesh.

Fig 3.23

It’s fair to say that during this stage of the mesh generation step, the user according to his knowledge of how the PANAIR code needs the geometry input, and in which manner he has decided to divide the aircraft will decide where each surface mesh needs to start. Keeping track of where each surface starts will make the translation of the grid points to the PANAIR format much simpler, and will save time.

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3.3.2 Modification of CATIA model for making automatic mesh possible It was clear that the mesh of the aircraft had to be done in a way that the points would be easy transferred to the PANAIR input file. The methods explaining the mesh are described in section 3.3.1 and these methods are followed throughout the parameterized mesh model. Going through every single meshed surface on the model would be too extensive so to describe the working path, some specific surfaces are discussed in this section. It has also been very important to mesh the parts in a specific order to be able to translate them to PANAIR for further analysis. This will be described more thoroughly in the section involving the Translation program. Basically every cut-out requires re-surfacing to be mesheble if one should follow the mesh rules described in previous sections. One example is the pylons for the engine under wing configurations.

Fig 3.24

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As can be seen in Fig 3.24 the pylons intersect the lower part of the mid wing, which has to be split into three surfaces so that a mesh based on the rules, discussed in the previous section, can be performed. The split surfaces on the wing are highlighted on Fig 3.25.

Fig 3.25

Both the wing and the engine have gone through modifications to be able to abide the mesh rules. So is every other surface on the model which happens to have a cut-out. In figure 3.26 the mesh of this section is in focus.

Fig 3.26

This part of the project came as a surprise as none expected that special treatment was necessary to a surface with a cut out. The reason for that being that the surface must always have exactly four sides, unless it is the beginning or the end of a body which then may have three sides, a so called collapsed surface. A great amount of time was spent to deal with this unexpected problem, which ultimately delayed the whole project.

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3.3.3 Parameterization of the meshed product Mesh as described in the previous section cannot be modified when the Model has changed configurations, as one will soon realize that the mesh becomes unstructured and inapplicable. The Mesh must be parameterized in such a way that it is referenced, by defining Radius/Length of the surface in question when creating the mesh. This will make the mesh part dependable on the references and it will make modifications possible.

Fig 3.27

The goal with the Meshed Aircraft Model, see Fig 3.27, has always been that it should be able to be automatically update the made mesh whenever the Parametric Aircraft Model is modified, so that the user doesn’t have to re-do the meshing process. Beside the fact that every surface needs special care to be able to be meshed correctly, one must always have the Translation program in mind so that it can correctly export the data to PANAIR. Therefore if done correctly, the Automatic/Parametric Meshed Aircraft parts will more or less guarantee that the data being transferred to PANAIR via the Translation program will be in a correct form.

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3.3.4 The Automatic Mesh Environment

Fig 3.28

Above picture illustrates how the mesh is parameterized. The user can at any time change the density of the nodes in each network, without having to re-do any of the mesh setups described in the section 3.3.1. In Fig 3.29, the network for the “cockpit” is highlighted and it shows that it is using the parameters “Elements_in_Radius” (=20) and “Cockpit” (=14), see Fig 3.28.

Fig 3.29

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3.3.5 Examples of different configurations The mesh parts are connected to the main Parametric Aircraft Model. Although being a separated part, these parts are still interconnected with the main model and therefore modified semi automatically whenever the main part is modified.

Fig 3.30

Due to certain limitations in the CATIA Analysis module, there was a need to create several mesh parts for different tail and engine configurations. See fig 3.31, Fig 3.30 and Fig 3.27.

Fig 3.31

The problem lies in the fact that some mesh surfaces still search for surfaces which have been de-activated in the main aircraft model. This will result in an error which essentially ruins the mesh model in question. Therefore for the time being different configurations have their own mesh part and all of these parts are linked to the main Aircraft model. One mesh part is the “one engine in back” configuration. This part can be modified for a wide range of aircraft from the Cessna Citation series to Embraer-145 to MD-90 as some examples. There are two other mesh parts, being “one engine under wing configuration” and “two engines under wing configuration”. To see examples of the automatic mesh, see Appendix 6.1.1.

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4 Translation Program

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4.1 Fundamentals To be able to analyze the aerodynamic specification of the aircraft model in PANAIR one must find a way to link CATIA with this particular panel code program. It was clear that after meshing the whole Aircraft one would be able to get the nodes from CATIA in a txt file. The problem lies in the fact that the output file of CATIA V5 has to be extensively modified. Therefore it is necessary to re-order the points in a way which is required by PANAIR and described in chapter 2. This is absolutely not a task to be done manually, though we are talking about thousands of points which need to be re-ordered. The Translation program which is made in Ada 95 has the following characteristic;

Re-ordering of the point defining the networks. Re-do the numbers in CATIA output file which are written with “,” as the decimal point,

unlike “.” used in PANAIR. Ex: 3,14 => 3.14. It was also necessary to make the user inputs semi-automatic as well, so that the translation

program asks the user questions like; “Input Mach number:” and “Input Angle of Attack:”. This is to make the PANAIR input file free from possible user mistakes.

The translation program also includes built in abutment characteristics. This in its essence means that all network edge points are modified (if necessary) and totally intersected before even going into PANAIR. Please see Appendix 6.2.2 for more information.

All other characteristics which should be in a PANAIR input file are made in the Translation program, such as proper “PANAIR abutments”, wakes etc. see Appendix 6.2.1 and 6.2.3.

Basically after running the Translation program one doesn’t need to do more than to run the input file, it gives, in PANAIR.

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4.2 Implementation

4.2.1 Environment As explained in the previous section, what the Translation program does is that it in its essence will work as a link between CATIA V5 and PanAir 502i.

Fig 4.1

The program searches for a file where some important reference parameters for PanAir are available. These parameters are exported directly from the Parametric Aircraft Model. Figure 4.1 shows both the parameters in CATIA and the same parameters exported to a txt file.

Fig 4.2

The points for the geometry of the Aircraft is gathered from a file called CATIA_Nodes, which has been exported from the Automatic Mesh Model. It is in this file the Translation program finds the points it needs to re-order and by doing so describe the geometry for PanAir. See figure 4.2.

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When one runs the Translation program, it starts to gathers the information from the two txt files described above, together with some user inputs and delivers the PanAir_input file.

Fig 4.3

Because of the fact that the automatic Mesh in CATIA never changes and only gets modified, there is no reason to believe that the Link will ever give an incorrect input file. However if the user as much as modifies a single line or point in the aircraft model, then this could change the characteristics of the Mesh and ultimately give a non working Link between PanAir and CATIA. It is however safe to say that everything will run smoothly as long as the user only modifies the aircraft by adjusting the parameters.

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4.2.2 Results An example of the input file given by the Translation program can be found in Appendix 6.2.3. The input file is that of the Cessna Citation configuration and is originally 50 pages long and has therefore been shortened by taking out all points besides the two first ones in each network. An exception is the first network where all points are preserved so that it gives the reader an idea of how the rest of the original input file layout looks like. Because of the sheer size of the output file for the full PAM configuration which amounts to almost 3000 pages (!), only one part which is considered to be of interest can be found in Appendix 7.2.4. There are two types of PAM input configurations. One is only with the fuselage and wing (see appendix 6.1.2) and another configuration is all components incorporated in the PAM, excluding the engine (see appendix 6.1.3). The following configurations have been exported and analyzed in PANAIR (see Appendix 6.1.1):

Cessna Citation Airbus 310 Boeing 777

Considering the fact that each PANAIR run takes about 30 min to perform, there have been only a limited, selected amount of testing. For all the tests that have been made, the Cl1 stands for the pressure integration method’s Cl and Cl2 stands for Trefftz method’s Cl. These methods are discussed in section 2.3.

Fig 4.4 Fig 4.5

One important issue that has been evaluated in chapter 2 is the importance of how to input the wing points. Fig 4.4 has been proven in the mentioned chapter to be the worse method of the two. When tested in the translation program, this method gave very bad results as well, while when inputting the points as Fig 4.5 has shown improved values and therefore all results presented in this section have been made following this pattern. The following results will be discussed in section 5.4. For the Cessna Citation configuration the following tests have been made:

Fuselage plus the wing: An Array of low angles of attack for Mach 0.5 Full configuration minus the engine: An array of low angles of attack for Mach 0.5 & 0.7

For the Airbus 310 configuration the following tests have been made:

Full configuration minus the engine: An array of low angles of attack for Mach 0.5 For the Boeing 777 configuration the following tests have been made:

Full configuration minus the engine: An array of low angles of attack for Mach 0.5

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Cessna Citation fuselage plus wing configuration

Mach 0.5

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

-4 -2 0 2 4 6 8

alpha

Cl

Cl1Cl2

Mach 0.5

0,00000,00500,01000,01500,0200

0,02500,03000,03500,04000,0450

-4 -2 0 2 4 6 8

alpha

Cdi

Fig 4.6

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Cessna Citation full configuration minus the engine

Mach 0.5

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

-4 -2 0 2 4 6 8

alpha

Cl

Cl1Cl2

Mach 0.5

0,00000,00500,01000,01500,02000,02500,03000,03500,04000,04500,0500

-4 -2 0 2 4 6 8

alpha

Cdi

Fig 4.7

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Mach 0.7

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

1,2000

-4 -2 0 2 4 6 8

alpha

Cl

Cl1Cl2

Mach 0.7

0,0000

0,0100

0,0200

0,0300

0,0400

0,0500

0,0600

0,0700

-4 -2 0 2 4 6 8

alpha

Cdi

Fig 4.8

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Airbus 310 full configuration minus the engine

Mach 0.5

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

1,2000

-4 -2 0 2 4 6 8

alpha

Cl

Cl1Cl2

Mach 0.5

0,0000

0,0100

0,0200

0,0300

0,0400

0,0500

0,0600

-4 -2 0 2 4 6 8

alpha

Cdi

Fig 4.9

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Boeing 777 full configuration minus the engine

Mach 0.5

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

-4 -2 0 2 4 6 8

alpha

Cl Cl1

Cl2

Mach 0.5

0,0000

0,0100

0,0200

0,0300

0,0400

0,0500

0,0600

-4 -3 -2 -1 0 1 2 3 4 5 6 7

alpha

Cdi

Fig 4.10

The results in this section will be discussed in chapter 5.4.

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4.2.3 Validation During the last days of this thesis work we were given some test results made by NASA (Reference 15) on a full-scale semi span Business Jet. A duplication of the given aircraft (see Fig. 4.11) was made in the PAM (see Fig. 4.12) and the geometry exported and analyzed in PANAIR 502i. Following are some specification and the test results on Mach 0.1 and 0.2. As can be seen in Fig 4.13 the airfoil was not duplicated correctly, only the thickness of the airfoil which in this case is 13% of the chord, simply because a mathematical formula of the airfoil in question couldn’t be found.

Fig 4.11 [15]

Fig 4.12

NASA PAM Semispan [m] 6,81 6,81Semispan Area [m] 11,61 11,61Root chord [m] 2,52 2,52Tip chord [m] 0,89 0,89Mean Chord 1,87 1,87Dihedral 5,50 5,50LE sweep 4,15 4,15Airfoil HSNLF 0213 NACA 2413 Fuselage Length [m] 10,69 10,69Fuselage Radius [m] 0,76 0,76Mach ~0,1 0,1 & 0,2

Fig 4.13

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Mach 0.1

-0,6000

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

-10 -5 0 5 10

alpha

Cl Cl1

Cl2

Mach 0.1

0,0000

0,0100

0,0200

0,0300

0,0400

0,0500

0,0600

0,0700

0,0800

-10 -5 0 5 10

alpha

Cdi

Fig 4.14

Mach 0.2

-0,6000

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

-10 -5 0 5 10

alpha

Cl Cl1

Cl2

Mach 0.2

0,0000

0,0100

0,0200

0,0300

0,0400

0,0500

0,0600

0,0700

0,0800

-10 -5 0 5 10

alpha

Cdi

Fig 4.15

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Comparison between M0.1 & M0.2

-0,6000

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

-10 -5 0 5 10

alpha

Cl Mach 01

Mach 02

Fig 4.16

Comparison between M0.1 & M0.2

0,0000

0,0100

0,0200

0,0300

0,0400

0,0500

0,0600

0,0700

0,0800

-10 -5 0 5 10

alpha

Cdi Mach 0.1

Mach 0.2

Fig 4.17

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5 Discussion and Conclusion

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5.1 The thesis work overall

Fig 5.1

The figure above shows the different stages that where achieved according to the goals that where set at the beginning of the project. The reader must remember that point (4) and (5) where set as “extra” goals and were to be tackled depending on the time remaining at the end of point (3). Clearly time was not enough because of the magnitude of the project, which at the beginning was not fully understood. However, not accomplishing point (4) does not mean that there is no information available gathered form this project. There have been some research by the authors, not presented in this report, that has shown the possibility to import arbitrary force gradients in to the CATIA FEM analyzes module and the little results that where gathered are promising. Therefore it is the conclusion of the authors that by investing time on continuing the line of research, finalizing points (4) and (5) should be possible. The following pages will give the conclusion for point (1) - (3).

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5.2 Parametric Model

In order to validate the Parametric Aircraft model, the Parametric Aircraft model was handed to be tested by the students of the course TMAL51 Aircraft Conceptual Design. The goal of their project was to design a Business Jet that could also be modified to an ambulance configuration. Overall the students could achieve the configuration they had in mind, but few of them wanted to go outside the boundaries of the parameters, hence the need to step in and modify the model for those specific students. The feedback was overall good, though they got configurations which they were more or less pleased with. With the feedback, some changes have been made on the model, but it will and should probably be more improvements in the future. One very important factor is that the model is relatively easy to modify and therefore upgrades will not be of any difficulties. Another important point is that the model did not break down during these trials, which indicates that the foundations are relatively solid. The first conclusion that can be drawn from the feedback given by the student’s, is that PAM does work in the range that was intended for, Jet civil aircraft, and it is robust enough. It’s important to remember that the model is very simplified so that it can be flexible enough to take the wide range of shapes and configurations it now is capable of taking. However it is not enough for Detail Design, hence the configurations the students have developed with the help of this model are not as detailed as they would have wished for. Being objective with the capabilities of the CAD model it is necessary to point out that the model is limited when it comes to the fairing. It is a complex part, surface wise. The Part is roughly constructed and limited to relatively few shapes and provides a window of opportunity to improve the model. Following is a list of pictures from the different configuration made in the mentioned course by some of the students. Note that surfaces such as windows, doors and winglets were added by the students themselves.

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Fig 5.2

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5.3 PANAIR Let us start the discussion of this section talking about the meshing tool used during this work. The meshing tool plays a key roll in this project since it is step to link the Parametric Model with the analysis software. When the project started and the CAD system was defined there was a certain amount of time spent looking for an appropriate tool that would perform the meshing of the surfaces. During the research it came to the attention of the authors that CATIA V5 has incorporated in the analysis environment an advance meshing tool, this made it simple to decide after a few test that it was a good meshing program having the advantage that since it was already incorporated to the CAD program. The biggest advantage is that it would not present any kind of problems with the surfaces. And at the end it prove to be a good meshing tool and easy to use. However, it is still a simple meshing tool and it is the opinion of the authors that it would be good for the project to research other programs that can be found in the market today. The relevance on having the best meshing tool possible lies on the fact that the quality of the mesh impacts directly on the quality of the results given by PANAIR. PANAIR 502i is a proven program in the industry, according to the research that has been conducted. There for it was a good option to implement for this methodology, this program is very sensitive to different aspects such as data format, geometry input, mesh density, etc. This made it difficult to validate the results during the project. Even so, for validating the surface distribution the methodology used in chapter 2.3 prove to be very useful in order to understand the logic of the program when talking about network input. From this surface validation it can be said that the proper way to divide the meshed surface is following the configuration of model SWB10. The reason to choose this configuration lies on the capability of the meshing tool, which here prove to be hard to adapt to the variation in the surface of the model, giving another reason why to search for a more advance meshing tool. Validations of the actual results given by the solver have been done and the results from these are promising as they will be shown further on this chapter (section 5.4). Only after these are properly finished then the second stage of a larger project can begin. This project is said to be an optimization tool for aircraft design that will comprehend the aerodynamic analysis given by this thesis work and incorporate a structural analysis module. The structural analysis module depends in great measure on the successful accomplishment of this work since from the results given by PANAIR the forces can be extracted and analyzed with an FEM software.

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5.4 Translation program

5.4.1 Comparison between different PAM configurations

Pressure

-0,4000

-0,20000,0000

0,20000,4000

0,6000

0,80001,0000

1,2000

-4 -2 0 2 4 6 8

alpha

Cl

Cl Boeing

Cl Cessna

Cl Airbus

Mach 0.5

0,0000

0,0100

0,0200

0,0300

0,0400

0,0500

0,0600

-4 -2 0 2 4 6 8

alpha

Cdi

Cdi Boeing

Cdi Cessna

Cdi Airbus

Fig 5.3

Overall what can be said from the test runs that have been made on the PAM, it can be said that the results given by PANAIR are with in a reasonable range. In Fig 5.3 the Cl comparison between the different configurations are calculated with the pressure integration method and the CDi is calculated with the Trefftz plane analysis method. The reason for this choice is explained in chapter 2. The tests are performed using a Mach of 0.5.

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5.4.2 Comparison between different Cessna surface configurations

Pressure

-0,4000

-0,2000

0,0000

0,2000

0,4000

0,6000

0,8000

1,0000

-4 -3 -2 -1 0 1 2 3 4 5 6 7

alpha

Cl Cl Full

Cl F&W

Mach 0.5

0,00000,00500,01000,01500,02000,02500,03000,03500,04000,04500,0500

-4 -3 -2 -1 0 1 2 3 4 5 6 7

alpha

Cdi Cdi Full

Cdi F&W

Fig 5.4

In fig 5.4 a comparison between different characteristics of Cl and CDi of a Cessna with different surface inputs, being one with only the fuselage and the wing (Cl- and CDi-F&W) and another being a full configuration without the engine (Cl- and CDi-full). See appendix 6.1.2 and 6.1.3. One can see that the when all of the Cessna components (excluding the engine) are exported to PANAIR, results in a higher Cl and CDi than the Cessna where only the fuselage and the wing are exported. The tests here are also performed using a Mach of 0.5. Any thorough and specific conclusion cannot be inputted, though the sources of error are yet too many and need to be narrowed done and this can only be done by a vast amount of testing and parameter changing.

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5.4.3 Validation Following is the comparison between the NASA Business Jet and the PAM model which was made so that it could have the basic characteristics of this Jet, as can be seen in Fig 4.12.

-0,8000

-0,7000

-0,6000

-0,5000

-0,4000

-0,3000

-0,2000

-0,1000

0,0000

0,1000

0,2000

0,3000

0,4000

0,5000

0,6000

0,7000

0,8000

0,9000

-10 -5 0 5 10

Fig 5.5

The comparison which is made in chapter 4.2.3, between Mach numbers 0.1 and 0.2, shows that the differences in Cl are marginal. However since the test results from NASA are more in the area of M 0.1, the results for M=0.1 of the PAM is considered to be more of interest and therefore compared. These comparisons can be seen in Fig 5.5 and 5.6.

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Fig 5.6

The result which is given by PANAIR doesn’t coincide fully with the wind tunnel test of NASA. The reasons for this are plenty. To start with the wing profiles are not the same, only thickness of the profiles are. The shape of PAM is not a 100% replica of the NASA Jet, though not enough data of the mentioned Jet could be found. Another possibility could be a not dense enough grid intensity of the meshed model. However considering the fact that the results are in the same scale is quite promising. One should however consider the fact this is the first and only comparison to wind tunnel data that has been made using this methodology. More comparisons have to be made for a final validation. If the method would prove consistency, one of many conclusions is making another Parametric Aircraft model for unconventional design, such as a BWB configuration, which would be a good complement for the statistical and empirical approaches which are limited to conventional designs.

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6 Appendix

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6.1 CATIA

6.1.1 Examples of different mesh configurations

Airbus-340 configuration

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Airbus-310 configuration

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Boeing-777 configuration

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Embraer-145 configuration

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Cessna Citation configuration

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6.1.2 Fuselage plus Wing configuration for PANAIR analyzes

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Fuselage Section

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Front fuselage + Cockpit

Middle fuselage

Tail Cone

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Wing section for both upper and lower surface

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6.1.3 Full Configuration minus the Engine for PANAIR analyzes

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Fuselage Section

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Front fuselage + Cockpit

Middle fuselage

Tail Cone

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Wing section for both upper and lower surface

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Horizontal Tail Section for both upper and lower surface

N23 & N24

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Vertical Tail Section

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6.1.4 Pre-set Engines Engine 1: Pratt & Whitney PW-4000-94=> Boeing 747, 767; MD-11, A-300, 310 Engine 2: Pratt & Whitney GP7000 => A-380 Engine 3: Rolls-Royce Trent 500 => A-340 Engine 4: Rolls-Royce Trent 1000 => B-787 Dream Liner Engine 5: General Electric CF6-80C2=> MD-11, Airbus A300, Boeing 767, 747-300 Engine 6: Honeywell TFE731 => Falcon 900, Learjet 36, Raytheon Hawker 800 Engine 7: Rolls-Royce Trent 800 => B-777 Engine 8: Rolls-Royce AE 3007=> Embraer-145 Engine 9: FJ44-4A => Cessna Citation CJ4 Engine 10: Pratt & Whitney F117-PW-100=> Boeing C-17

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6.2 Translation program

6.2.1 Edge Numbering

Fuselage Section

Cockpit

Front fuselage

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Middle fuselage

Tail Cone

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Wing section for both lower and upper surface

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Horizontal Tail Section for both lower and upper surface

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Vertical Tail Section

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Wing wakes Section

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6.2.2 Forced Intersections

Artificial intersections are needed to be done by the Translation program between Networks:

• 17, 18 and 19 • 16, 17 and 18 • 16 and 33 • 17 and 34 • 19 and 35

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Artificial intersections are needed to be done by the Translation program between Networks:

• 5, 6 and 7 • 26, 7, 9 and 11 • 27, 6, 8 and 10 • 26 and 28 • 27 and 29 • 10, 11 and 12

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6.2.3 Input File of the “Full Configuration minus the Engine”

$title Parametric Aircraft Model mehta $datacheck 0. $symmetry - xz plane of symmetry =misymm mjsymm 1. 0. $mach number =amach 0.5 $cases - no. of solutions =nacase 3. $angles-of-attack =alpc 4. =alpha(1) alpha(2) alpha(3) 0. 2. 4. $printout options =isings igeomp isingp icontp ibconp iedgep 0. 0. 0. 1. 1. 0. =ipraic nexdgn ioutpr ifmcpr 0. 0. 1. 0. $references for accumulated forces and moments =xref yref zref nref 0. 0. 0. =sref bref cref dref 23220952. 8580. 2347. 13790. $points - wing-body with composite panels =kn cpnorm 20. 2. =kt 1. =nm nn netname 8. 11. body1 =x(1,1) y(1,1) z(1,1) x(*,*) y(*,*) z(*,*) 0.00 0.00 -400.00 18.10 0.00 -474.40 63.39 0.00 -536.61 122.31 0.00 -586.45 187.74 0.00 -627.48 256.63 0.00 -662.44 327.63 0.00 -692.92 400.00 0.00 -720.00 0.00 0.00 -400.00 16.98 10.07 -471.46 61.26 19.09 -532.98 120.05 26.80 -582.85 185.93 33.52 -623.98 255.39 39.50 -658.79 327.03 44.97 -689.12 400.00 50.05 -716.01 0.00 0.00 -400.00 16.06 20.12 -467.29 59.27 38.37 -526.59 118.11 53.95 -575.29 184.24 67.23 -615.28 254.32 78.86 -649.10 326.50 89.31 -678.40 400.00 98.86 -704.27 0.00 0.00 -400.00 15.35 29.85 -461.92 57.84 57.10 -517.57 116.62 80.56 -563.57 183.01 100.08 -601.33 253.49 116.93 -633.24 326.09 131.83 -660.80 400.00 145.24 -685.06 0.00 0.00 -400.00 14.86 39.05 -455.25 56.94 75.02 -505.82 115.68 105.39 -547.97 182.24 130.94 -582.24 252.97 152.49 -611.42 325.84 171.31 -636.64 400.00 188.06 -658.85 0.00 0.00 -400.00 14.72 47.63 -447.58 56.64 91.39 -491.58 115.37 128.15 -528.43 181.99 158.73 -558.43 252.80 184.16 -584.41 325.76 206.49 -606.65 400.00 226.27 -626.27 0.00 0.00 -400.00 14.86 55.25 -439.05 56.94 105.80 -475.04 115.68 147.97 -505.39 182.24 182.25 -530.93 252.97 211.41 -552.51

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325.84 236.63 -571.33 400.00 258.85 -588.06 0.00 0.00 -400.00 15.35 61.92 -429.85 57.84 117.56 -457.11 116.62 163.57 -480.56 183.01 201.34 -500.07 253.49 233.23 -516.94 326.09 260.79 -531.84 400.00 285.06 -545.24 0.00 0.00 -400.00 16.06 67.29 -420.12 59.26 126.58 -438.38 118.11 175.29 -453.95 184.24 215.28 -467.22 254.32 249.09 -478.88 326.50 278.39 -489.33 400.00 304.27 -498.86 0.00 0.00 -400.00 16.98 71.46 -410.07 61.26 132.98 -419.10 120.05 182.85 -426.81 185.93 223.98 -433.52 255.39 258.79 -439.51 327.03 289.12 -444.98 400.00 316.01 -450.05 0.00 0.00 -400.00 18.07 74.41 -400.00 63.42 136.59 -400.00 122.31 186.44 -400.00 187.72 227.51 -400.00 256.64 262.41 -400.00 327.64 292.88 -400.00 400.00 320.00 -400.00 8. 11. body2 0.00 0.00 -400.00 18.07 74.41 -400.00 21. 15. body3 400.00 0.00 -80.00 400.00 50.05 -83.95 21. 15. body4 3200.00 0.00 800.00 3200.00 125.14 790.13 21. 8. body5 5847.50 0.00 800.00 5847.50 125.14 790.14 7. 2. body6 6433.37 806.62 -572.07 6433.37 726.89 -733.18 15. 2. body7 6433.37 0.00 800.00 6433.37 125.81 790.03 7. 15. body8 6465.25 804.40 -588.16 6465.25 735.74 -725.84 15. 15. body9 6465.25 0.00 800.00 6465.25 123.88 790.34 7. 11. body10 7700.75 773.98 -671.31 7700.75 680.80 -775.23 15. 11. body11 7700.75 0.00 800.00 7700.75 118.31 791.18 21. 2. body12 8862.92 0.00 800.00 8862.92 127.55 789.73 21. 8. body13 8936.25 0.00 800.00 8936.25 138.33 787.92 21. 4. body14 9554.00 0.00 800.00 9554.00 125.14 790.13 21. 6. body15 10083.50 0.00 800.00 10083.50 125.14 790.09 21. 5. body16 10950.01 0.00 800.00 10950.01 119.62 790.54 8. 10. body17 11612.27 97.97 792.98 11612.27 198.52 770.68 14. 10. body18 11612.27 638.97 366.33 11612.27 669.77 268.04 21. 6. body19 12893.97 44.73 797.49 12893.97 136.94 775.86 21. 4. body20 13436.14 0.00 800.00 13436.14 34.34 797.26 $points - bodybase with composite panels =kn cpnorm 2. =kt 5. 6. 11. body21 =x(1,1) y(1,1) z(1,1) x(*,*) y(*,*) z(*,*) 13790.00 0.00 720.00 13790.00 0.00 736.00 6. 11. body22 13790.00 0.00 720.00 13790.00 16.00 720.00 $points - bodybase with composite panels Horizontal Tail =kn cpnorm 3.

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=kt 1. 10. 15. body23 =x(1,1) y(1,1) z(1,1) x(*,*) y(*,*) z(*,*) 12893.97 393.18 320.05 12755.06 440.60 331.53 10. 15. body24 11612.27 655.53 320.00 11745.26 650.92 276.98 10. 2. body25 13365.17 2900.00 320.41 13286.30 2900.00 321.68 $points - bodybase with composite panels Wing Parts =kn cpnorm 7. =kt 1. 26. 2. body26 =x(1,1) y(1,1) z(1,1) x(*,*) y(*,*) z(*,*) 8862.92 812.36 -535.00 8747.15 813.52 -518.24 26. 2. body27 6433.37 806.62 -572.07 6465.25 804.40 -588.16 26. 15. body28 8833.78 879.99 -535.00 8738.08 880.00 -521.74 26. 15. body29 6465.24 879.99 -535.00 6530.18 880.00 -599.22 26. 26. body30 8984.36 3880.03 -431.25 8915.33 3879.54 -417.21 26. 26. body31 7269.09 3880.00 -430.23 7318.58 3881.49 -473.05 26. 2. body32 8528.45 8579.99 -266.11 8546.10 8579.45 -250.35 $points - bodybase with composite panels Vertical Tail =kn cpnorm 4. =kt 1. 19. 5. body33 =x(1,1) y(1,1) z(1,1) x(*,*) y(*,*) z(*,*) 13238.18 0.00 2720.00 13116.18 0.00 2606.70 19. 10. body34 13552.46 53.35 2720.00 13444.65 54.98 2612.92 19. 6. body35 14365.57 33.77 2720.00 14283.81 34.33 2613.19 19. 2. body36 13238.18 0.00 2720.00 13309.57 31.56 2720.00 $points - body to wing wakes =kn cpnorm 1. =kt 19. =nm nn netname 27. 2. awbw =x(1,1) y(1,1) z(1,1) x(*,*) y(*,*) z(*,*) 8862.92 812.36 -535.00 8936.25 810.51 -410.13 $trailing wakes from wings 3. =kt 18. body26 1. 22000. .0 wake1 body28 1. 22000. .0 wake2 body30 1. 22000. .0 wake3 $trailing wakes from Horizontal Tail 1. =kt 18. body24 3. 22000. .0 wake4 $points - body wake 2. =kt matcw 18. 1.

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13. 2. wake5 =x(1,1) y(1,1) z(1,1) x(*,*) y(*,*) z(*,*) 13790.00 0.00 800.00 13790.00 12.51 799.00 9. 2. wake6 13790.00 76.04 695.29 13790.00 71.25 683.69 $pea - partial or full network edge abutments =nfpa iopfor movusr 38. 0. 2. =nne peatol 2. =nn en epinit eplast body1 2. body2 4. 3. body3 4. body2 3. body1 3. 2. body3 2. body4 4. 2. body4 2. body5 4. 3. body5 2. body6 4. body7 4. 2. body6 2. body8 4. 2. body8 2. body10 4. 2. body7 2. body9 4. 2. body9 2. body11 4 3. body12 4. body11 2. body10 2. 2. body12 2. body13 4. 2. body13 2. body14 4. 2. body14 2. body15 4. 2. body15 2. body16 4. 3. body16 2. body17 4. body18 4. 3. body19 4. body17 2. body18 2. 2. body19 2. body20 4. 2. body23 4.

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body17 3. 2. body24 4. body18 1. 2. body23 3. body24 1. 2. body23 1. body24 3. 4. body26 4. body7 3. body9 3. body11 3. 2. body26 2. body28 4. 2. body28 2. body30 4. 4. body27 4. body6 1. body8 1. body10 1. 2. body27 2. body29 4. 2. body29 2. body31 4. 2. body26 1. body27 3. 2. body26 3. body27 1. 2. body28 3. body29 1. 2. body28 1. body29 3. 2. body30 3. body31 1. 2. body30 1. body31 3. 2. body33 3. body16 1. 2. body34 3. body17 1. 2. body35 3. body19 1. 2. body33 2. body34 4. 2. body34 2. body35 4. 2. awbw 2. wake1 4.

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2. wake1 2. wake2 4. 2. wake2 2. wake3 4. 2. wake1 1. body26 1. 2. wake2 1. body28 1. 2. wake3 1. body30 1. $eat - liberalized abutments 0. 0. 0. 0. 1. $TREFFTZ PLANE ANALYSIS

$end of a502 inputs

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6.2.4 Output File of the “Full Configuration minus the Engine” ----------------------------------------------------- full configuration forces and moments summary ----------------------------------------------------- symmetry conditions: misymm = 1 mjsymm = 0 sol-no alpha beta cl cdi cy ------ ------- ------- ------- ------- --------- --------- --------- --------- ------------ 1 2.0000 0.0000 0.51881 1.25266 0.00000 Trefftz plane analysis: cl = 0.366431E+00 cdi = 0.865792E-02 eff = 0.155714E+01 2 4.0000 0.0000 0.73315 1.28981 0.00000 Trefftz plane analysis: cl = 0.599348E+00 cdi = 0.233141E-01 eff = 0.154702E+01 3 6.0000 0.0000 0.93977 1.33587 0.00000 Trefftz plane analysis: cl = 0.831530E+00 cdi = 0.458487E-01 eff = 0.151420E+01 ******************************************************************************

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6.3 Surface Division Validation Results

6.3.1 Lift Coefficient results (Cl)

ALPHA Surface Pressure Treffzt Surface Pressure Treffzt 2 Cl diff (%) Cl diff (%) Cl diff (%) Cl diff (%)

SWB 0.08159 0.08214 C-SWB 0.08993 0.09335 SWB1 0.08059 1.23 0.0805 2.01 C-SWB1 0.08792 2.24 0.09072 2.82 SWB2 0.08044 1.41 0.0815 0.76 C-SWB2 0.08913 0.89 0.09254 0.87 SWB3 0.14467 -77.31 0.11997 -46.06 C-SWB3 0.13546 -50.63 0.11272 -20.74 SWB4 0.08177 -0.22 0.0821 0.00 C-SWB4 0.09026 -0.37 0.0933 0.01 SWB5 0.04854 40.51 0.03378 58.88 C-SWB5 0.05323 40.81 0.03429 63.27 SWB6 0.08106 0.65 0.07991 2.71 C-SWB6 0.08157 9.30 0.08330 10.77 SWB7 0.03774 53.74 -0.23252 383.08 C-SWB7 0.05488 38.97 0.03414 63.43 SWB8 0.04095 49.81 0.07533 8.30 C-SWB8 0.08164 9.22 0.08330 10.77 SWB9 0.04005 50.91 0.07487 8.85 C-SWB9 0.08093 10.01 0.08259 11.53

SWB10 0.08109 0.61 0.08226 -0.15 C-SWB10 0.08085 10.10 0.08259 11.53 4

SWB 0.1628 0.1639 C-SWB 0.17155 0.1759 SWB1 0.16061 1.35 0.1605 2.09 C-SWB1 0.16780 2.19 0.17097 2.79 SWB2 0.16091 1.16 0.1630 0.57 C-SWB2 0.17012 0.83 0.17468 0.69 SWB3 0.19429 -19.34 0.17869 -9.02 C-SWB3 0.15812 7.83 0.15803 10.15 SWB4 0.16311 -0.19 0.1639 0.00 C-SWB4 0.17206 -0.30 0.1759 0.00 SWB5 0.09507 41.60 0.06695 59.15 C-SWB5 0.10533 38.60 0.06821 61.22 SWB6 0.15947 2.05 0.15874 3.16 C-SWB6 0.16356 4.66 0.16602 5.61 SWB7 0.09053 44.39 -0.19058 216.27 C-SWB7 0.10652 37.91 0.06806 61.31 SWB8 0.11992 26.34 0.15415 5.96 C-SWB8 0.16363 4.62 0.16602 5.61 SWB9 0.11821 27.39 0.15340 6.41 C-SWB9 0.16233 5.37 0.16496 6.21

SWB10 0.16225 0.34 0.16436 -0.27 C-SWB10 0.16226 5.42 0.16496 6.21 8

SWB 0.32224 0.3267 C-SWB 0.33205 0.3401 SWB1 0.31769 1.41 0.3197 2.12 C-SWB1 0.32490 2.15 0.33069 2.78 SWB2 0.31964 0.81 0.3252 0.47 C-SWB2 0.33006 0.60 0.33816 0.58 SWB3 0.85952 -166.73 0.29529 9.61 C-SWB3 0.66748 -101.02 0.24769 27.18 SWB4 0.32276 -0.16 0.3267 0.00 C-SWB4 0.33283 -0.23 0.3402 0.00 SWB5 0.18713 41.93 0.13302 59.28 C-SWB5 0.20780 37.42 0.13575 60.09 SWB6 0.31348 2.72 0.31571 3.36 C-SWB6 0.32479 2.19 0.33068 2.78 SWB7 0.19571 39.27 -0.10499 132.14 C-SWB7 0.20799 37.36 0.13561 60.13 SWB8 0.27542 14.53 0.31112 4.76 C-SWB8 0.32481 2.18 0.33068 2.78 SWB9 0.27283 15.33 3.10E-01 5.17 C-SWB9 0.32305 2.71 0.3289 3.30

SWB10 0.32228 -0.01 3.28E-01 -0.33 C-SWB10 0.323 2.73 0.3289 3.30

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6.3.2 Induced Drag Coefficient Results (CDi)

ALPHA Surface Pressure Treffzt Surface Pressure Treffzt 2 CDi diff (%) CDi diff (%) CDi diff (%) CDi diff (%)

SWB 0.00080 0.0013 C-SWB 0.00198 0.00165 SWB1 0.00114 -42.50 0.0012 3.72 C-SWB1 0.00235 -18.69 0.00156 5.14 SWB2 0.00432 -440.00 0.0013 1.42 C-SWB2 0.00511 -158.08 0.00162 1.38 SWB3

0.00943 -

1078.75 0.0028 -114.68 C-SWB3

0.01243 -527.78 0.00247 -49.88 SWB4 0.00089 -11.25 0.0013 0.00 C-SWB4 0.0021 -6.06 0.0016 0.02 SWB5 -0.04365 5556.25 0.0007 44.11 C-SWB5 -0.05999 3129.80 0.00074 55.12 SWB6 -0.03942 5027.50 0.0012 5.19 C-SWB6 -0.00036 118.18 0.00131 20.40 SWB7 -0.01395 1843.75 0.2230 -17281.7 C-SWB7 -0.05224 2738.38 0.00073 55.51 SWB8 -0.04188 5335.00 0.0011 15.77 C-SWB8 -0.00036 118.18 0.00131 20.40 SWB9 -0.03776 4820.00 0.0011 16.75 C-SWB9 0.00275 -38.89 0.0013 21.63

SWB10 0.002 -150.00 0.0013 -0.25 C-SWB10 0.00274 -38.38 0.0013 21.64 4

SWB 0.00682 0.0051 C-SWB 0.0079 0.0058 SWB1 0.00707 -3.67 0.0049 3.87 C-SWB1 0.00811 -2.66 0.00552 5.06 SWB2 0.01042 -52.79 0.0051 1.06 C-SWB2 0.01127 -42.66 0.00575 1.18 SWB3 -0.06186 1007.04 0.0061 -18.56 C-SWB3 -0.04769 703.67 0.00539 7.29 SWB4 0.00692 -1.47 0.0051 0.00 C-SWB4 0.00805 -1.90 0.0058 0.01 SWB5 -0.04096 700.59 0.0028 44.81 C-SWB5 -0.05605 809.49 0.00293 49.72 SWB6 -0.03378 595.31 0.0048 6.00 C-SWB6 0.00533 32.53 0.00520 10.59 SWB7 -0.01200 275.95 0.2167 -4146.17 C-SWB7 -0.04835 712.03 0.00291 49.94 SWB8 -0.03767 652.35 0.0045 11.37 C-SWB8 0.00534 32.41 0.00520 10.59 SWB9 -0.03348 590.91 0.0045 12.19 C-SWB9 0.00868 -9.87 0.0051 11.63

SWB10 0.00817 -19.79 0.0051 -0.49 C-SWB10 0.00868 -9.87 0.0051 11.63 8

SWB 0.03055 0.0203 C-SWB 0.03121 0.0218 SWB1 0.03043 0.39 0.0195 3.94 C-SWB1 0.03082 1.25 0.02067 5.06 SWB2 0.0342 -11.95 0.0201 0.88 C-SWB2 0.03500 -12.14 0.02154 1.05 SWB3 -0.37520 1328.15 0.0166 18.24 C-SWB3 -0.36763 1277.92 0.01561 28.27 SWB4 0.03069 -0.46 0.0203 0.00 C-SWB4 0.03143 -0.70 0.0218 0.00 SWB5 -0.02841 193.00 0.0111 45.14 C-SWB5 -0.03997 228.07 0.01159 46.76 SWB6 -0.01128 136.92 0.0190 6.38 C-SWB6 0.02824 9.52 0.02063 5.23 SWB7 0.00099 96.76 0.2057 -914.68 C-SWB7 -0.03256 204.33 0.01156 46.88 SWB8 -0.01789 158.56 0.0184 9.08 C-SWB8 0.02825 9.48 0.02063 5.23 SWB9 -0.01367 144.75 0.0183 9.82 C-SWB9 0.03201 -2.56 0.0204 6.15

SWB10 0.03223 -5.50 0.0204 -0.61 C-SWB10 0.03202 -2.60 0.0204 6.15

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7 References

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